EP1308611A2 - Cloison pare-feu pour turbines à gaz - Google Patents

Cloison pare-feu pour turbines à gaz Download PDF

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Publication number
EP1308611A2
EP1308611A2 EP02257234A EP02257234A EP1308611A2 EP 1308611 A2 EP1308611 A2 EP 1308611A2 EP 02257234 A EP02257234 A EP 02257234A EP 02257234 A EP02257234 A EP 02257234A EP 1308611 A2 EP1308611 A2 EP 1308611A2
Authority
EP
European Patent Office
Prior art keywords
zone
gas turbine
engine
turbine engine
fire
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Granted
Application number
EP02257234A
Other languages
German (de)
English (en)
Other versions
EP1308611B1 (fr
EP1308611A3 (fr
Inventor
Olivier Jean-François Cazenave
Robert Edward Sherlock
Tim Andrew Clark
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Rolls Royce PLC
Original Assignee
Rolls Royce PLC
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by Rolls Royce PLC filed Critical Rolls Royce PLC
Publication of EP1308611A2 publication Critical patent/EP1308611A2/fr
Publication of EP1308611A3 publication Critical patent/EP1308611A3/fr
Application granted granted Critical
Publication of EP1308611B1 publication Critical patent/EP1308611B1/fr
Anticipated expiration legal-status Critical
Expired - Lifetime legal-status Critical Current

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Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D9/00—Stators
    • F01D9/06—Fluid supply conduits to nozzles or the like
    • F01D9/065—Fluid supply or removal conduits traversing the working fluid flow, e.g. for lubrication-, cooling-, or sealing fluids
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
    • F02C7/24—Heat or noise insulation
    • F02C7/25—Fire protection or prevention
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
    • F02C7/32—Arrangement, mounting, or driving, of auxiliaries
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B22—CASTING; POWDER METALLURGY
    • B22F—WORKING METALLIC POWDER; MANUFACTURE OF ARTICLES FROM METALLIC POWDER; MAKING METALLIC POWDER; APPARATUS OR DEVICES SPECIALLY ADAPTED FOR METALLIC POWDER
    • B22F2998/00—Supplementary information concerning processes or compositions relating to powder metallurgy

Definitions

  • the invention relates to gas turbine engines for aircraft and particularly to the routing of pipes and harnesses within such engines.
  • Gas turbine engines are divided into a number of 'fire zones', the different fire zones tending to operate at respectively different temperatures when the engine is functioning, and being separated by 'fire walls'.
  • the fire walls prevent any flammable fluid leakage between the various zones and help to prevent the spread of a fire from one zone to another.
  • the core region of the engine comprises one or two zones, normally referred to as Zones 2 and 3, and the region outside the fan case constitutes a separate zone, referred to as Zone 1.
  • Zones 2 and 3 the region outside the fan case constitutes a separate zone, referred to as Zone 1.
  • Zone 1 There is also a bypass zone in the region in which bypass air flows, between the core and the fan case.
  • Zone 3 is located generally radially inwards of Zone 1. However, typically Zone 3 is extended radially outwardly and downwardly into the general area of Zone 1, for a limited circumferential extent, in a lower region of the engine. This extended Zone 3 region forms a "bifurcation", because bypass air is forced to pass around it, the air being directed by a splitter fairing.
  • the above described prior art arrangement has certain disadvantages.
  • the pipes and harnesses are usually required to cross the fire wall at a right-angle.
  • These constraints affect the pipe and harness routing, creating unnecessary bends.
  • the resulting intricate design makes clashes more likely and harder to detect during design and makes modifications from an original design more difficult to accommodate and unnecessarily complex.
  • the routing of the pipes and harnesses appears untidy and is therefore difficult to follow and likely to cause confusion and reduce accessibility for maintenance. It is also necessary to provide large numbers and various different types of brackets, lugs and clips to hold the pipes and harnesses. This increases the manufacturing costs and increases the removal/refit time.
  • a further disadvantage of the prior art arrangement relates to the routing of the radial drive.
  • a D-seal between a radial drive shroud and the bifurcation disconnect panel.
  • an O-ring seal provided between the radial drive shroud and the transfer gearbox has been known to fail in service, producing oil leakage.
  • the size of the D-seal support dictates how closely the splitter fairing can be wrapped around the radial drive. A larger splitter fairing is less aerodynamically efficient than a small one.
  • a gas turbine engine including:
  • the member forms a fire wall between the second fire zone and the bifurcation part of the first fire zone.
  • the member is mounted to extend downwardly from the fan case, when the engine is in its normal orientation as mounted on an aircraft.
  • the member may be mounted on a rear part of the fan case.
  • the member may include side portions which are preferably planar and which are preferably angled at between 30° and 60° to the axial direction of the engine. Most preferably, the side portions are angled at between 40° and 50° to the axial direction of the engine . The side portions preferably each lie in a respective generally vertical plane.
  • the member may further include a mid-portion which joins the side portions.
  • the mid-portion may also be planar and is preferably oriented at about 90° to the axial direction of the engine. Preferably the mid-portion also lies in a generally vertical plane.
  • the member may further include a mounting portion which may be adapted for attachment to the fan case.
  • the mounting portion may extend forward from the mid-portion at a top of the member and may lie in a plane which is generally perpendicular to the plane of the mid-portion.
  • the mounting portion may be slightly curved and may be of a complementary shape to the fan case.
  • the member may comprise a sheet material which may be steel or titanium.
  • the thickness of the material is at least 1mm and most preferably the thickness is at least 0.4mm.
  • the gas turbine engine may further include a radial drive connecting a turbine drive shaft to a gearbox mounted on the fan case.
  • the radial drive passes through the fan case.
  • a bellow-seal is provided between the radial drive and the fan case.
  • the bellow-seal may be secured on to the radial drive shroud by a jubilee clip.
  • the bellow-seal may be connected to the fan case through a bolted flange.
  • the bellow-seal may be provided with a ring having a spigot providing a hard surface for bolting to the fan case.
  • the seal allows the radial drive to move relative to the fan case.
  • the seal is also fireproof.
  • a ducted fan gas turbine engine generally indicated at 10 comprises, in axial flow series, an air intake 12, a propulsive fan 14, an intermediate pressure compressor 16, a high pressure compressor 18, combustion equipment 20, a high pressure turbine 22, an intermediate pressure turbine 24, a low pressure turbine 26 and an exhaust nozzle 28.
  • the gas turbine engine 10 works in the conventional manner so that air entering the intake 12 is accelerated by the fan 14 to produce two air flows, a first air flow into the intermediate pressure compressor 16 and a second airflow of "bypass air" which provides propulsive thrust.
  • the bypass air flows within a nacelle 33 which surrounds part of the engine including the propulsive fan 14.
  • the nacelle 33 is defined by radially inner and outer walls 35, 37 respectively.
  • the intermediate pressure compressor 16 compresses the air flow directed into it before delivering the air to the high pressure compressor 18 where further compression takes place.
  • the compressed air exhausted from the high pressure compressor 18 is directed into the combustion equipment 20 where it is mixed with fuel and the mixture combusted.
  • the resultant hot combustion products then expand through and thereby drive the high, intermediate and low pressure turbines 22, 24 and 26 before being exhausted through the nozzle 28 to provide additional propulsive thrust.
  • the high, intermediate and low pressure turbines 22, 24 and 26 respectively drive the high and intermediate pressure compressors 16 and 18 and the fan 14 by suitable interconnecting shafts.
  • Fig. 2 illustrates the gas turbine engine 10 in somewhat more detail. It may be seen that the engine 10 includes a fan case 32 which defines an outer boundary of a bypass zone through which the bypass air passes. Generally externally of the fan case 32 and between the walls 35, 37 there is defined a further fire zone, Zone 1 and near to the core of the engine there are defined two further, hotter zones, Zones 2 and 3.
  • Each zone is separated from the adjacent zones by a fire wall.
  • the fire walls prevent flammable fluid leaking between the zones and help prevent the spread of a fire starting in one of the zones.
  • Zone 3 is generally located radially inwardly of Zone 1. However, for a circumferentially limited extent at the base of the engine, Zone 3 is extended radially outwardly into Zone 1.
  • the cross section of Fig. 2 has been taken approximately vertically through a centre of the engine and therefore through this radially extended Zone 3 region.
  • a splitter fairing 42 divides and deflects the bypass air such that it passes smoothly around the downwardly extending Zone 3 region.
  • Fig. 3 illustrates the disconnect panel 46 in more detail. It may be seen that the pipes and harnesses 44 cross the disconnect panel 46 substantially perpendicularly thereto. There is also a requirement for a minimum straight section before and after the disconnect region where the pipes/harnesses 44 cross the panel, as well as a minimum bend radius. This results in an intricate and complex design, as may be seen from Fig. 3.
  • a radial drive 48 also passes through the bifurcation disconnect panel 46.
  • the radial drive 48 transfers power from the high pressure turbine shaft to a gearbox 49 mounted on the fan case 32 in Zone 1.
  • a D-seal 50 is provided between the radial drive 48 and the bifurcation disconnect panel 46.
  • the seal 50 may be generally P-shaped or any other shape as known in the art.
  • a member in the form of a disconnect panel 52 which replaces the bifurcation disconnect panel 46 of the prior art.
  • the disconnect panel 52 is mounted on the fan case 32 so as to extend downwardly and radially outwardly therefrom.
  • the disconnect panel 52 includes a mounting portion 53 which is adapted to be mounted on the fan case 32, the mounting portion being either generally planar or curved in a complementary way with the fan case 32. Extending downwardly from the mounting portion is a generally planar centre portion 56 which is approximately rectangular. Extending from each outer edge of the centre portion 56 is a side portion 58, each side portion 58 also being generally planar and rectangular. The side portions 58 are oriented at an angle of approximately 40° to 50° to the axial direction of the engine. Each portion of the disconnect panel 52 is made from a sheet material such as steel or titanium. The thickness of the sheet material is about 0.4 to 0.5mm. The disconnect panel 52 is able to withstand 1100°c for 15 minutes with a standard flame producing 116k W/m 2 ⁇ 10kW/m 2 .
  • the disconnect panel 52 is so positioned and shaped that the side portions 58 lie across a natural route for most of the pipes and harnesses 44 passing from Zone 3 to Zone 1 and the centre portion 56 lies on a natural route for the remainder of the pipes and harnesses 44.
  • the disconnect panel 52 forms a fire wall between Zones 1 and 3 and also forms a mounting and a disconnect means for the pipes and harnesses 44. Because the side and centre portions 58 and 56 lie across a natural route for the pipes and harnesses 44, routing is straightforward and it is easy to ensure that the pipes cross the disconnect panel 52 at the correct angle of approximately 90°.
  • disconnect panel 52 therefore eliminates prior art constraints on routing, allowing a tidy layout to be achieved during design. This eases design and re-design, clashes are more likely to be avoided and access to and identification of pipes and harnesses 44 is more simple. There is also the scope to improve the clipping of pipes and harnesses 44, as clamp blocks can be accommodated. Furthermore, one type of clamp can be used in several locations since all the pipes are parallel. This reduces the amount of additional clips used, which results in shorter assembly and removal time and lower pipe costs.
  • the radial drive 48 may pass through and be sealed against the fan case 32 near to the lower end of the radial drive, where efficient sealing is easier to achieve.
  • the radial drive 48 may be sealed with a bellow seal 60 which enables the radial drive to move freely relative to the fan case 32. Vibration from excitation to the splitter fairing 42 will therefore not be fed to the radial drive 48.
  • the bellow seal 60 does not prevent free displacement of the radial drive 48 locally, any difficulties with an O-ring seal provided between the radial drive 48 and the transfer gear-box disappear.
  • the bellow seal 60 allows a reduction of the gap between the splitter fairing 42 and the radial drive 48 which improves the aerodynamics of the engine.

Landscapes

  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Physics & Mathematics (AREA)
  • Fluid Mechanics (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)
EP02257234A 2001-11-02 2002-10-17 Cloison pare-feu pour turbines à gaz Expired - Lifetime EP1308611B1 (fr)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
GBGB0126371.4A GB0126371D0 (en) 2001-11-02 2001-11-02 Gas turbine engines
GB0126371 2001-11-02

Publications (3)

Publication Number Publication Date
EP1308611A2 true EP1308611A2 (fr) 2003-05-07
EP1308611A3 EP1308611A3 (fr) 2009-09-30
EP1308611B1 EP1308611B1 (fr) 2010-12-08

Family

ID=9925051

Family Applications (1)

Application Number Title Priority Date Filing Date
EP02257234A Expired - Lifetime EP1308611B1 (fr) 2001-11-02 2002-10-17 Cloison pare-feu pour turbines à gaz

Country Status (4)

Country Link
US (1) US20030126854A1 (fr)
EP (1) EP1308611B1 (fr)
DE (1) DE60238517D1 (fr)
GB (1) GB0126371D0 (fr)

Cited By (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB2497934A (en) * 2011-12-22 2013-07-03 Rolls Royce Plc A gas turbine aeroengine arrangement
EP3217001A1 (fr) * 2016-03-07 2017-09-13 United Technologies Corporation Moyeu de paroi pare-feu
EP3239497A1 (fr) * 2016-04-25 2017-11-01 United Technologies Corporation Montage de module électronique sur une structure d'isolation de vibrations
EP3677751A1 (fr) * 2019-01-04 2020-07-08 Rolls-Royce plc Bifurcation de moteur à turbine à gaz

Families Citing this family (9)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US7010906B2 (en) * 2001-11-02 2006-03-14 Rolls-Royce Plc Gas turbine engine haveing a disconnect panel for routing pipes and harnesses between a first and a second zone
DE102004023569A1 (de) * 2004-05-13 2005-12-08 Mtu Aero Engines Gmbh Flugtriebwerk
US20060032974A1 (en) * 2004-08-16 2006-02-16 Honeywell International Inc. Modular installation kit for auxiliary power unit
US9284887B2 (en) 2009-12-31 2016-03-15 Rolls-Royce North American Technologies, Inc. Gas turbine engine and frame
EP2719621A3 (fr) * 2012-10-10 2017-10-25 Sikorsky Aircraft Corporation Système d'échappement arrière pour aéronef à voilure tournante
US9482157B2 (en) 2013-02-28 2016-11-01 United Technologies Corporation Bifurcation fire purge system for a gas turbine engine
US10156188B2 (en) * 2016-04-25 2018-12-18 United Technologies Corporation Electronic module location for mechanical components
GB201716499D0 (en) * 2017-10-09 2017-11-22 Rolls Royce Plc Gas turbine engine fireproofing
US11118705B2 (en) 2018-08-07 2021-09-14 General Electric Company Quick connect firewall seal for firewall

Citations (1)

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Publication number Priority date Publication date Assignee Title
US20050274485A1 (en) 2004-06-14 2005-12-15 Huggins George L Cast unitized primary truss structure and method

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US4437627A (en) * 1982-03-12 1984-03-20 The Boeing Company Integrated power plant installation system
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US5174110A (en) * 1991-10-17 1992-12-29 United Technologies Corporation Utility conduit enclosure for turbine engine
EP0694120B1 (fr) * 1993-03-03 2001-05-16 KETEMA AEROSPACE & ELECTRONICS DIVISION Systeme integre de regulation pour turbomoteur

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Publication number Priority date Publication date Assignee Title
US20050274485A1 (en) 2004-06-14 2005-12-15 Huggins George L Cast unitized primary truss structure and method

Non-Patent Citations (1)

* Cited by examiner, † Cited by third party
Title
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Cited By (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB2497934A (en) * 2011-12-22 2013-07-03 Rolls Royce Plc A gas turbine aeroengine arrangement
GB2497934B (en) * 2011-12-22 2014-06-04 Rolls Royce Plc Aeroengine arrangement
EP3217001A1 (fr) * 2016-03-07 2017-09-13 United Technologies Corporation Moyeu de paroi pare-feu
US10526915B2 (en) 2016-03-07 2020-01-07 United Technologies Corporation Firewall mount hub
EP3239497A1 (fr) * 2016-04-25 2017-11-01 United Technologies Corporation Montage de module électronique sur une structure d'isolation de vibrations
US10047676B2 (en) 2016-04-25 2018-08-14 United Technologies Corporation Electronic module mounting to vibration isolating structure
EP3677751A1 (fr) * 2019-01-04 2020-07-08 Rolls-Royce plc Bifurcation de moteur à turbine à gaz

Also Published As

Publication number Publication date
US20030126854A1 (en) 2003-07-10
GB0126371D0 (en) 2002-01-02
EP1308611B1 (fr) 2010-12-08
EP1308611A3 (fr) 2009-09-30
DE60238517D1 (de) 2011-01-20

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